The first rotating turbine stage in an axial flow gas turbine engine is subject to the harshest combination of environmental factors, including surface temperature, materials stress, etc. The turbine blades, disposed about the periphery of the rotor disk, interact with the hot engine working gas at temperatures commonly in excess of 2000.degree. F. (1100.degree. C.), while the rotor disk itself is subject to high radial loading induced by the rapid angular velocity of the spinning stage.
It is common in modern high performance gas turbine engines to protect the turbine blades with relatively cool compressed air. This flow of air is received by the blade at the radially inward attachment portion and distributed throughout by internal airflow passages, exiting the individual blade through a series of small openings at the blade surface or from the root portion.
The compressed air is provided by the upstream compressor section of the gas turbine and is available at the turbine disk in an axially flowing, annular stream. This annular stream, having inwardly bypassed the combustor section, is radially inward of the turbine blades and must therefore be redirected and distributed outward. This distribution is commonly effected through a manifold formed by securing an axially spaced faceplate to the turbine disk thus providing a radial passage for the annular compressed airstream to the blades. It will be appreciated by those skilled in the art that the addition of a faceplate is preferable to providing air flow paths within the highly loaded rotor disk, avoiding stress concentrations which may in turn reduce disk strength or induce premature disk cracking.
The securing of the faceplate to the disk must involve some type of cooperative mechanical engagement therebetween. The use of bolts or other fasteners which require perforating the disk and/or faceplate is to be avoided to the extent possible for the reasons just noted. One technique known in the prior art for axially securing at least a portion of the faceplate to the disk is by the use of a plurality of corresponding hooked protrusions or dogs disposed in both the faceplate and the rotor disk.
During assembly, the disk and faceplate are placed into axial contact with the hooked protrusions of the faceplate disposed intermediate the corresponding hooked protrusions of the disk, the disk and faceplate subsequently rotated each with respect to the other for aligning and engaging the corresponding hooked protrusions. This engagement opens up radial flow passages intermediate the engaged dogs, allowing a free flow of cooling air to the turbine blades.
The faceplate and disk remain thus engaged so long as relative rotation between the disk and faceplate does not occur. Even a slight relative rotational displacement, not sufficient to disengage the dogs, may block off a portion of the radial airflow paths leading to a reduction of cooling airflow and eventual blade overheating. Prior art methods for preventing such relative rotational displacement include the use of bolts or locating pins inserted between the faceplate and rotor following assembly. Prior art manifold systems and anti-rotation structures are well disclosed in U.S. Pat. Nos. 3,010,696 and 4,435,123 issued respectively to Everett and Levine.
The prior art methods, while effective, require holes to be provided in the disk and faceplate, leading to undesirable local stresses in the corresponding member. In addition, the presence of bolts or locating pins along with their associated retaining structure increases both the number of individual parts and complexity of the turbine rotor stage as well as the likelihood that the faceplate and/or disk may be scratched or damaged during assembly.